Missile configurations,controls and utilization techniques

ABSTRACT

Missile configurations with engine and propellant control systems are the subject of this application together with weapons and navigational techniques employing same. Reaction engine control systems employing relatively moveable plug-cowl configurations with associated control systems are described herein for providing control of thrust direction and magnitude, engine operating conditions, missile kinematics, and other parameters of liquid and solid propellant rockets.

United States Patent Parilla [S4] MISSILE CONFIGURATIONS,

CONTROLS AND UTILIZATION TECHNIQUES [72] Inventor: Arthur R. Parilla,PO. Box 127,

Mountain Lakes, NJ. 07046 [22] Filed: Oct. 3, 1968 [21] Appl. No.:767,583

Related US. Application Data [62] Division of Ser. No. 607,068, Jan. 3,1967, Pat.

[52] US; Cl. ..244/3.21, 60/230, 60/242 [5]] Int. Cl ..F42b 15/18, F02k1/08, F02k 1/16 [58] Field of Search ..244/l, 3.21, 3.22

[56] References Cited UNITED STATES PATENTS 2,584,127 2/1952 Harcum eta1. ..244/3.2

[451 Sept. 19,1972

3,210,937 10/1965 Perry, Jr. ..60I35.6 X

Primary Examiner-Benjamin A. Borchelt Assistant ExaminerThomas B. WebbAttorney-Morgan, Finnegan, Durham and Pine [57] ABSTRACT Missileconfigurations with engine and propellant control systems are thesubject of this application together with weapons and navigationaltechniques employing same; Reaction engine control systems employingrelatively moveable plug-cowl configurations with associated controlsystems are described herein for providing control of thrust directionand magnitude, engine operating conditions, missile kinematics, andother parameters of liquid and solid propellant rockets.

31 Claims, 12 Drawing Figures PATENTED 3.692.258

SHEET 2 0F 4 TIMING f 551%? I ACCEL' error cowL-pwa VELOC. ACTUATORGEOME my MISS/LE ssusma 1 LR I TERM REFESRENCE J, l R GRAM 0AM I L l L 5ae a 1 52 51 J I vswclrv I CONTROL 1 v-ACTUA Ton ACCELER.

CONTROL FIG-.6

INVENTOR.

ARTHUR R. PAR/LLA A T TOR/V635 PATENTEDSEP 19 I972 3.692.258

sum 3 or 4 GUIDAME A T T! T00! CDNTRDL VEC TOR CONTROL FIG.8

SENSED DATA: /0

'MISSILE KINEMA TICS osmorv ENGINE x=am=mm1=i'sm's I PARAMETER l3 SENSORcQMP A AmR ERROR 051'. I2 REFEOQENCE 1 "ME PROGRAM FACTORS AC u T0 Z6 1T A R INVENTOR.

: ARTHUR R. PARILLA FEREN BY or CE (L404 1/2.

ROGRAMMED DATA F l 9 ATTOfiA/EKS'.

PATENTEDSEP 19 1 12 3.692.258

saw u 0F 4 :3 FIG. 12 A ,4 TTORNE Y5 MISSILE CONFIGURATIONS, CONTROLSAND UTILIZATION TECHNIQUES RELATED PATENTS AND APPLICATIONS Thisapplication is a division of applicants prior copending application,Ser. No. 607,068, filed Jan. 3, 1967 for Missile Configurations,Controls and Utilization Techniques now US. Pat. No. 3,489,373 which isa continuation-in-part of applicants prior copending application, Ser.No. 302,222, filed June 14, 1963 and now abandoned, for AircraftMissiles, Missile Weapons System and Space Craft." Another division ofapplicants prior copending application, Ser. No. 607,068 copending withthis application is application, Ser. No. 784,818, filed Dec. 18, 1968,for Solid Propellant Rocket Engine Control And Missile Configurations.

The latter is in turn a copending division of application, Ser. No.701,571, filed Dec. 9, 1957 and now US. Pat. No. 3,094,072, granted June18, 1963. An additional application of the applicant, Ser. No. 860,304,is also related to the instant application, being anothercontinuation-in-part of Ser. No. 701,571, and now US. Pat. No.3,151,446.

This invention relates to control systems and operating techniques formissiles having engines capable of controllable thrust magnitude anddirection. Examples of such engines are found in applicants above citedpatents.

Described herein are apparatus and techniques for controlling variousflight and engine parameters of such missiles to provide improvedaccuracy, flexibility and reliability of operation. lnstrumentalitiesand techniques are disclosed for controlling thrust magnitude anddirection, missile velocity, acceleration and attitude, engineefficiency and thrust termination over a wide range of environmentalkinematic and aerodynamic conditions.

According to the invention new concepts in antispace missile weaponssystems are provided. Defensive missiles are described for interceptingintercontinental ballistic missiles, satellites, or other space objectsat extreme altitudes to provide defense over large areas. In order toinsure accuracy of interception, the defensive missile according to theinvention is capable of hovering flight at zero velocity in space forfinite time periods while the guidance system seeks and locks onto thetarget, followed by a collision course in any direction to the targetupon command of the guidance system. The defensive missile is capable ofmaneuverability in space in any direction in any of the three principalplanes, covering a range of several hundred miles from its positionduring hovering flight. This flexibility may be provided with eithersolid or liquid propellant rocket engines. Applicants aforementioned US.Pat. No. 3,489,373 contains a detailed description of the implementationof the control systems and methods of the present application to achievehovering flight and antispace missile weapons systems.

To achieve the foregoing, novel control system arrangements are providedin which various flight parameters such as acceleration, velocity,incremental velocities, or their combination, may be maintained constantindependent of variable drag due to variable altitude, or otherenvironmental conditions, and independent of variations in missile mass,engine perforrnance due to manufacturing tolerances, and othervariables.

These new operational and control techniques permit improvedanti-ballistic missile weapons systems for local defense againstballistic missile attack; improved ground-to-air defense; and improvedlong range air-tosurface missiles of minimum weight and maximumreliability.

BACKGROUND Most jet engines in which thrust is developed by thedischarge of a compressible fluid at supersonic velocities, such asrocket engines, ram-jet engines or turbojet engines, or combinations ofthese such as the ramjet ducted-rocket, or the turbo-rocket, usuallyemploy nozzles of the converging-diverging type, sometimes called aDeLaval or venturi nozzle, characterized by a fixed geometry having aconstant throat area and constant ratio of area expansion in thediverging portion.

It is well recognized that a variable throat area nozzle improves theflexibility of engine operation over a broader range; thrust may bevaried in magnitude as the throat area is increased or decreased; thecombustion characteristics and internal gas dynamics upstream of thethroat being greatly affected by the magnitude of the throat area.

Engine efficiency is also improved by a variable nozzle expansion ratio,resulting in increased thrust or engine efficiency. The desired areaexpansion ratio is a function of pressure ratio or ratio of internalchamber pressure at nozzle inlet to ambient atmospheric pressure atnozzle discharge. Even in engines which operate at substantiallyconstant chamber pressure, as in liquid rocket engines, the pressureratio increases with altitude. If the nozzle is designed for sea-levelconditions, under-expansion and loss of efficiency occurs at higheraltitudes. If it is designed for a high altitude condition,over-expansion occurs at sea-level resulting in a loss of thrust duringthe critical take-off condition.

Most engines employing fixed geometry nozzles, are designed to operateat fixed thrust levels. While some degree of throttleability is possiblein liquid propellant rocket engines and in air breathing engines bymeans of control valves which reduce mass flow rate of propellant orfuel to the engine, this range is limited. Combustion chamber pressurereduces at the lower mass flow rates with a constant nozzle throat area,further throttling being limited by combustion instability or flame-outat the lower chamber pressures. The range of throttleability maytherefore be increased by a variable throat area which maintains chamberpressure within operable limits even at reduced thrust.

This lack of flexibility, or throttleability is even greater in solidpropellant rocket engines, for once ignition occurs, further controlover thrust is impossible in conventional engines. The thrust timecharacteristic is dependent upon the internal ballistics, i.e., thepropellant burning rate, propellant burning surface, propellant densityand nozzle throat area, the ratio of propellant burning surface tonozzle throat area being an important parameter, denoted by the symbolK. Since these quantities are all fixed in the design stage, furthercontrol during flight is not possible. A variable nozzle throat areacontrollable during flight can, then, by varying K exert an importantcontrol over thrust.

In a similar manner, control over thrust termination is a seriousproblem with solid propellant engines. While this is accomplishedreadily in liquid propellant rocket engines and in air breathingengines, simply by valve closure, conventional solid propellant rocketengines require special provisions to provide controlled andreproducible thrust tennination, which is critical for accurate controlof ballistic missiles. One such method is by ejection of a nozzleinsert, the abrupt increase in nozzle throat area causing a rapidpressure drop which extinguishes combustion. It is found thatundesirable thrust peaks occur with this method as a result of thesimultaneous product of high chamber pressure and large throat area atthe moment ejection occurs. This is objectionable since the highaccelerations thus transmitted to the missile may damage sensitivecomponents.

Alternate means for thrust termination provide auxiliary nozzles ororifices which discharge combustion gases in a forward direction whenburst discs are energized, thus neutralizing the thrust from the primarynozzles discharging rearwardly. This creates a packaging problem fordisposition of the gases from the missile, and for loading propellant,while the additional components add weight and complexity to the system.Also, neither solution permits re-starts for those applications whereintermittent thrust may be desired in a controllable solid propellantrocket engine.

In applying propulsion systems to guided missiles, it is frequentlydesired to vary not only the magnitude of thrust, but also itsdirection. Means for thrust vector directional control can providemissile stability and maneuverability during launch and during flight.Substantial reduction in cost, weight and aerodynamic drag can beachieved by eliminating large aerodynamic fins, and complex fittingsrequired for fin attachment. This, then, reduces storage space andimproves accuracy by reducing dispersion due to aerodynamic gusts.

Present methods for thrust vector control generally require subsidiarymeans attached to or near the noule exit so as to deflect the jet streamwhen actuated. Examples are jet vanes, as used on the German V-2 rocket,and more recently, jetavators.

Jet vanes are submerged in the jet stream issuing from the nozzle,rotation of the vane causing deflection of the jet stream therebyvarying the direction of thrust. Since such vanes are constantlyimmersed in the jet stream, they cause drag losses reducing net thrusteven when control forces are not required. This objection is eliminatedin jetavators in which a circular ring mounted on gimbal supportssurrounds the jet stream at the nozzle exit. The ring is normally freeof the jet stream, but dips into it when actuated to deflect the jet andhence the resultant direction of thrust. Obviously, the jetavator mustbe of larger diameter than the nozzle exit, thereby limiting expansionratio, or requiring diameters larger than maximum diameter of rocketcase, in which event, aerodynamic drag replaces drag due to immersion injet stream. Large deflection angles are required to be effective, sinceonly a portion of the ring intersects the jet stream, the oppositediameter moving away from it. In both jet vanes and jetavators,additional components are required which must be constructed of hightemperature resistant materials, and which add weight, complexity andcost, and reduce reliability.

An alternate method of thrust vector control is used with liquidpropellant rocket engines in which the entire thrust chamber assembly ispivotally mounted on gimbal rings to provide freedom of motion in twoperpendicular planes. Obviously, the gimbal ring and mounting provisionsmust be designed to transmit the full thrust loads, adding weight. Also,control forces must be large to provide the desired high frequency ofresponse. A further problem exists because of the need for flexible feedlines for delivering propellant to the thrust chamber. As engine sizeincreases, these become large diameter pipes with the requirement forflexibility becoming increasingly difficult. This method is whollyimpractical for solid propellant rocket engines wherein the entire massof propellant is stored within the thrust chamber, resulting in a veryhigh moment of inertia.

Other problems associated with solid propellant rocket engines which maybe alleviated by the variable nozzle throat area technique described inthe aforementioned patents and in further detail hereinafter include:(l) the temperature sensitivity of the propellant to ambient temperatureof the grain before firing; (2) erosive burning, whereby burning rateincreases upon ignition because of high velocity gases passing overpropellant burning surfaces; (3) variation of propellant burning surfaceduring burning, including effects of small cracks in propellant; and (4)sliver fonnation, or small sections of unbumt propellant remaining afterburnout.

items (I), (2) and (3) result in peak chamber pressures, increasingdesign requirements, thereby adding weight, while (4) adds weight whilecontributing nothing to performance.

In addition to the foregoing, a major problem which is aggravated by thelimitations of present propulsion systems, is the successfulinterception of ballistic missiles, satellites or other objectstravelling at great velocities through or from outer space. The splitsecond precision required to insure collision and mutual destructionbetween two bodies travelling at great velocities is enormous.

These, and other problems hereinafter described, offer seriouslimitations to performance of aircraft, missiles, missile weaponssystems, and space ships.

it is, therefore, the purpose of this invention to advance the state ofthe art in missile configuration, control and utilization to overcomethe aforementioned limitations and to accomplish, in addition to thoseobjectives recited in applicant's U.S. Pat. Nos. 3,094,072 and 3,489,373the following specific objectives:

To provide mechanical or fluid spring control arrangements forautomatically controlling the variable throat area of jet enginenozzles.

To provide fluid, and/or electrically operated control systems andactuators for controlling the variable throat area of jet enginenozzles.

To provide improved ballistic missiles the flight trajectory of which isrelated to engine performance so as to provide optimum nozzle thrustcoefficient relationships as a function of altitude by automaticallyvarying the nozzle area expansion ratio.

To provide a missile system capable of maintaining a constantacceleration during flight independent of variable drag at variablealtitude, variable mass, or other environmental condition.

To provide a missile system capable of increasing its own velocity by afixed increment independent of variable drag at variable altitude,variable mass or other environmental condition.

To provide a missile system capable of maintaining a constant flightvelocity independent of variable drag at variable altitude, variablemass, or other environmental condition.

To provide a missile system capable of maintaining constant accelerationfollowed by constant velocity, or vice versa, and repetitively in anyprescribed manner as desired, independent of variable drag, variablemass, or other environmental condition.

To provide an anti-missile missile system capable of hovering flightwhile the guidance system seeks and locks onto the invading enemytarget, followed by high maneuverability of the defensive missile in anydirection as it accelerates on a collision course for targetinterception.

To provide a simple mechanical analog computer as part of the guidanceand control system of an anti-missile missile for directing thedefensive missile from its hovering position to the desired collisioncourse for target interception.

To improve the throttleability of liquid propellant rocket engines bymaintaining thrust chamber pressure at reduced propellant flow ratesautomatically by means of a novel variable area nozzle control system.

To provide new liquid propellant booster rockets incorporating newconcepts in thrust chamber control, expellant bag design for gaspressurization of propellant tanks, and improved reproducibility inpressurization by gas generators.

To provide greater safety by rendering missiles nonpropulsive duringstorage in the event of accidental ignition.

To protect rocket casings against destructive failure in the event ofmoderate grain cracking by use of a variable area nozzle acting as apressure relief valve.

Serving to illustrate exemplary embodiments of the invention are thedrawings of which:

FIG. I is a fragmentary elevational view, partially diagrammatic, andpartially sectional, of a hydraulic missile control system;

FIG. 2 is a view corresponding to FIG. 1, but employin g an electricalmissile control system;

FIG. 3 is a fragmentary elevational view, partially diagrammatic andpartially sectional, of a control system employing fluid spring means;

FIG. 4 is a sectional detail view of the variable fluid springs of thesystem of FIG. 3;

FIG. 5 is a block diagram illustrating a control system for controllingmissile kinematics;

FIG. 6 is a block diagram illustrating missile velocity and accelerationcontrol modes;

FIG. 7 is a partially diagrammatic, partially sectional view of ananalog computer employed in the direction control system of a missile;

FIG. 8 is a partially diagrammatic, partially sectional fragmentary viewof a portion of the computer of FIG.

FIG. 9 is a diagrammatic view of a general control arrangement forcontrolling missile performance;

FIG. 10 is an elevational view, partially in section, of a liquidpropellant rocket and control system therefor;

FIG. 11 is an elevational view, partially in section, of a liquidpropellant rocket and its control system embodying an expellant bag andnozzle control structure;

FIG. 12 is a fragmentary elevational view, partially in section,illustrating a modification t0 the system of FIG. I l;

A. GENERAL FEATURES OF VARIABLE COWL- PLUG CONFIGURATION l Control ofPressure, Thrust Magnitude, and Related Factors In applicants previouslycited prior copending applications and patents, techniques areillustrated for varying the throat area and other geometric parametersof nozzles employed in thrust producing devices. Among the illustrativeexamples are a number of embodiments for varying the throat area andstream orientation in rocket engine nozzles to achieve variations inexpansion ratio, chamber pressure, pressure ratio, propellant specificimpulse, thrust magnitude, thrust direction, and related parameters. Abrief review of the foregoing will serve as an aid to understanding theinvention.

FIG. 1 of US. Pat. No. 3,489,373 illustrates the implementation of thesetechniques in a plug-type rocket engine nozzle. Reference may be made toUS. Pat. No. 3,094,072 for further description of its structure andoperation.

In the nozzle of FIG. 1, the resultant pressure forces acting on thecowl 28 are transmitted by the rod 54, adjustable nut, 58, andcompression spring, 57, supported by the brackets 55 and 56 mounted onthe case 26.

The control forces to position the cowl may be supplied in any desiredmanner. The simple mechanical springs in FIG. 1 may be replaced byelectrical, hydraulic or pneumatic actuators, with signals from theguidance or control system to vary the nozzle position in any prescribedmanner, as described hereinafter.

2. Thrust Termination When adjusted by any of these means, the throatarea may be increased continuously, with a continuous decrease inchamber pressure, thereby achieving thrust tennination while avoidingthe abrupt discontinuity which occurs with other techniques such asnozzle insert ejection. In this manner, thrust decay can be reproduciblycontrolled simply by extending the cowl 28 so as to increase the throatarea to many times its normal design value. With propellant propertieswhich permit continued burning at low ambient atmospheric pressures, thecowl may again be retracted until, at super-critical pressure ratios,the thermal energy of the gases is again converted into kinetic energyof the jet stream, with high velocities directed aft, again renderingthe unit propulsive.

The above system also has the advantage that precision control over bothmissile velocity and attitude may be provided by operating at reducedthrust levels, with vector control, as described more fully hereinafter,available before thrust termination. Other means for thrust terminationare described in the above-noted US. Pat. No. 3,094,072 and illustratedin FIG. 2 of the aforementioned US. Pat. No. 3,489,373, where ring 52may be released by destruction of the explosive bolts 50.

3. Thrust Vector Direction Control In addition to thrust magnitudecontrol and termination, FIG. 2 of U.S. Pat. No. 3,489,373 alsoillustrates means for controlling the direction of the thrust vector. Asillustrated, the means involve mechanisms, e.g., linear actuators 64,for causing angular motion of the cowl about a transverse axis relativeto the cylinder 26 to deflect the stream. Further description is foundin applicants U.S. Pat. No. 3,094,072.

An alternate method for providing both translatory and oscillatorymotion of the cow] may be provided without gimbal rings by use of fouractuators 203 mounted 90 apart, as shown in FIG. 3 of the U.S. Pat. No.3,489,373 and described in greater detail in the U.S. Pat. No.3,094,072.

4. Combined Variable Area, Thrust Vector Control and Thrust Terminationwith Internal Supersonic Expansion The improvements shown herein are notnecessarily limited to nozzles with external expansion. The controlmechanisms may be employed where variable area, thrust vector controland thrust termination are combined within a nozzle employing onlyinternal expansion. This is shown in FIG. 4 of the U.S. Pat. No.3,489,373 and described in greater detail in said U.S. Pat. No.3,094,072.

B. PARAMETER REGULATING TECHNIQUES A variable throat area nozzle offersimportant mechanical solution to many problems in solid propellantrocket engines. It makes possible substantial reduction in case weightbeyond the use of higher strength materials by automatically maintainingconstant chamber pressure independent of the temperature sensitivity ofthe propellant, and of the progressivity or re gressivity of thepropellant grain.

It also makes possible a controllable solid propellant rocket enginewhose thrust may be varied at will, providing flexibility in operationwhich even surpasses throttleability of liquid propellant rocketengines.

FIG. 5 of the U.S. Pat. No. 3,489,373 illustrates the application of thevariable nozzle principles to a solid propellant rocket engine, which isdescribed in greater detail in the U.S. Pat. No. 3,094,072.

It may be seen that when the internal pressure increases (due to any ofseveral reasons as described below), the higher pressure force causesfurther compression of the spring 57, opening the cowl to a largerthroat area. Similarly, when the internal pressure decreases, the lowerpressure force causes the spring 57 to extend, thereby retracting thecowl, resulting in a smaller throat area. The throat area thus increasesautomatically with increasing chamber pressure.

Other parameter regulating techniques are described in the aforesaidpatent and include (I) constant pressure operation independent ofambient temperature (see FIG. 6 of the U.S. Pat. No. 3,439,373) (2)constant pressure operation independent of grain regressivity andprogressivity, (3) reduced thrust variation, and (4) adjustable thrustlevel.

PARAMETER REGULATING TECHNIQUES WITH EXTERNAL PROGRAMMABLE CONTROL Theuse of a variable area nozzle makes possible a controllable solidpropellant rocket engine in which the thrust may be varied in magnitudeat will throughout flight. Thrust variation may be obtained by varyingthe nozzle cowl position and, hence, throat area responsive to any typecontrol system.

1. General Control Techniques The mechanical springs for positioning thecowl as shown in FIG. 5 of the U.S. Pat. No. 3,489,373 may be replacedby servo-controls, either fluid or electrically actuated, illustrated inFIGS. 1 and 2 respectively.

In the systems shown in FIGS. 1 and 2, the difference between rocketchamber pressure and a control pressure from any source is used as theinput signal to increase or decrease thrust.

A system using fluid actuation is illustrated in FIG. I, while FIG. 2illustrates a system with electrical actuators. Further description ofthese systems and their uses in controlling thrust and thrusttermination are found in the U.S. Pat. No. 3,094,072.

A fluid spring installation is illustrated in FIGS. 3 and 4 in which thefluid spring assembly 159 replaces the mechanical spring 57 andassociated parts earlier described. As described in the U.S. Pat. No.3,094,072, this arrangement may be utilized for non-propulsive storage,thrust termination, and booster disposal.

2. Control Techniques for Regulating Acceleration and Velocity As notedabove, the adjustable cowl-plug configuration lends itself withconsiderable facility to control over various missile parameters by wayof self regulation and in response to external controls.

Control of thrust magnitude, thrust direction, chamber pressure, thrusttermination and related conditions such as those associated withnon-propulsive storage and booster disposal, has been describedhereinbefore.

These techniques are also applicable to the control of other parametersdealing with missile kinematics.

The accuracy and, hence, reliability with which missiles reach theirtargets may be greatly improved if the normal variations from round toround could be eliminated. These variations include missile weight,propellant loading, and engine performance due to manufacturingtolerances; environmental conditions, such as ambient temperatures andaltitudes; and other factors which vary the missile weight, drag andthrust relationships during flight.

As one example, in the case of air-to-air or airto-surface missiles, thedrag may vary considerably dependent upon the altitude of the carrierplane from which the missile is launched. Thus, the time to reach thetarget could vary due to cumulative variations. There are, of course,other examples such as an Anti-ICBM missile where great accuracy isrequired to insure successful interception of the attacking ICBM.

If the thrust of the variable nozzle rocket engine is controlledresponsive directly to some slight parameter, such as acceleration,velocity, or both, its accuracy could be greatly improved, cancellingnormal variations such as those set forth above.

In applicants above cited U.S. Pat. No. 3,094,072 systems are describedfor controlling the missile parameters acceleration and velocity.

A schematic indication of these systems for the case of accelerationcontrol is illustrated in FIG. 5. An accelerometer is oriented to sensemissile acceleration. Any appropriate acceleration measuring instru'ment may be employed, two types mentioned in said patent beingmass-spring units with either mechanical or fluid springs.

Desired acceleration, either a constant reference value, or programmablevalues for adjustable control during flight, is set into the system bymeans 171 which may be for example arrangements which deflect themechanical spring or pressurize the fluid spring.

The accelerometer output, indicative of actual missile acceleration iscoupled to an error detector 172 along with the desired accelerationdata.

The error detector may be of any known configuration and, as describedin the above-cited US. Pat. No. 3,094,072, can take the form of adifferentially actuated electrical switch arm or a difierentiallyactuated fluid valve spool.

Any discrepancy between actual and desired acceleration is reflected inan error indication manifested as an electrical or hydraulic signal.This signal is fed to the actuator 173 which causes a corrective changein the cowl-plug configuration.

When the correction is completed, actual and desired missileacceleration are brought into conformity.

The period of controlled acceleration may be adjusted by suitable timingmeans such as 175 which are adapted to disable the servo controlledacceleration mode.

For example, missile acceleration can be maintained for a controlledtime interval and then terminated by timing means 175 thus providing acontrolled velocity increment in missile performance. By regulatingacceleration as hereinbefore described during the time interval, adesired terminal velocity is thereby achieved. The timing device can atthis time then program a thrust termination into the power actuator.

Velocity control may be accomplished by substituting velocity sensingmeans in the system of FIG. thus providing an alternate method forestablishing desired missile velocity. Any velocity responsive means maybe used, such as those which sense the ratio of dynamic to staticpressure when the missile operates within the atmosphere or any othermeans.

Joint control is also obtainable in the manner shown in FIG. 6 by timedswitching of the cowl-plug actuator between a velocity control mode andan acceleration control mode. The timing may be programmed with localdevices such as motor-operated switches or from the guidance system.

3. Angle of Attack Vector Control System In general, the attitudesensing equipment in the missile, with or without external commands,supplies the required data to the thrust vector control systemhereinbefore described. In certain cases, such as one hereinafterdescribed, it is desired that the missile follow a course angle 0 whilemaintaining a missile heading B which is different from the courseangle. Under these conditions, the missile will fly at an angle ofattack (a=6-B), it being noted that outside the atmosphere lift and dragare negligible and therefore not relevant in missile movement.

A simple mechanical analog computer, shown in FIGS. 7 and 8, may be usedto quickly determine the proper missile heading, 3, for any desiredflight direction, 6.

The computer assembly, 300, of FIGS. 7 and 8 is designed so that aninput signal of angle 9 produces an output signal equal to angle B.

As shown in FIG. 7, the worm and wheel, 301, rotates the driver arm 302through the angle 0 by means of the shaft 303 containing a socket 304 inwhich the driver arm 302 is free to reciprocate. The driver arm rotatesthe pin 305 within the grooves 306 in the housings 307 and 308, thegrooves being concentric with the computer centerline 309. The driverarm 302 through pin 305 also rotates the driven arm 310 which is mountedin the socket 311, integral with the shaft 312, located on thecenterline 309. Thus, as shown in FIG. 8, rotation of the input shaft303 through angle 0, causes rotation of the output shaft 312 through theangle B. The guidance system then determines only the desired flightdirection, 0, the analog computer then supplying the correspondingmissile heading angle, B, to the vector control system.

The driver and driven arms 302 and 310 are both in the vertical planeduring climb and while hovering. When the guidance system signals adesired flight path of +0", plus being in the clockwise direction) thedriver shaft 303 rotates the arm 305 through 0, (counterclockwise), thedriven arm 310 and output shaft 312 rotating through B. Rotation ofshaft 312 introduces a bias, or error" in the control system whichmaintains the missile centerline coincident with the vertical referenceaxis through the Earth's center. The missile is thus rotated through theangle B, or until the driven arm 310 again coincides with the verticalreference axis, O-A. The missile will now have the desired heading, [3,to accomplish the flight path 0. The new missile heading will bemaintained constant with reference to the vertical reference axis by thesame nozzle vector control system as in vertical flight.

4. Generalized Control Techniques FIG. 9 graphically summarizes andgeneralizes the control techniques hereinbefore described forcontrolling missile behavior. In the illustration, the missile isidentified at 26 and includes adjustable cowl 28. The cowl isuni-directionally moved for the control of thrust magnitude anddifferentially moved for thrust vector direction control. Movement isaccomplished by the actuators 10 which may be any of the typeshereinbefore mentioned. The actuators 10 are driven by the error orcorrection signal derived from the comparator 11 which receives in turntwo inputs. One input comprises a signal representing the particularparameter which is being sensed, this being derived from the parametersensor 13. The other input is a command signal or signal indicative ofthe reference condition, derived from control means 12. If these signalsdo not correspond, indicating the actual condition differs from thedesired condition, an error or correction signal is developed. Thiserror signal operates the actuators 10 to bring the actual conditioninto correspondence with the desired or commanded condition.

Maintenance of the missile in a desired angular orientation relative toa reference line such as the vertical, was described hereinbefore. Insuch an application, the control device 12 is supplied with the desiredangular orientation data, e. g., the angle #0 With appropriatetranslation, this signal is fed to the comparator 11 which compares thissignal with the signal from sensor 13 which indicates the actual angularorientation of the missile. The sensor may be a vertical seekinggyroscope, or a stabilized platform or the like. If the command angle isdifl'erent from the actual angle sensed by the sensor, an error orcorrection signal is generated in the comparator 11 and fed to theactuators 10. In this application, the actuators are differentiallydriven to produce cowl rotation thereby causing the thrust vectordirection to change. In response, the missile commences to changeheading and when the actual heading corresponds with the desiredheading, the error signal disappears and the thrust vector controlsystem is accordingly deenergized after the usual stabilizing transient.For angle-of-attack types of control, the input device 12 may includethe to B converter shown in FIG. 7.

As noted hereinbefore and described more fully hereinafter, the commandsignal or reference level may be programmed into the system initially ormay be received from a remote point during flight. For example, desiredthrust levels may be preprogrammed in the control system with timingmeans employed to command the desired levels in accordance with theprogram. In the missile weapon system described hereinafter, desired orcommanded missile conditions are received in one mode from a remotepoint.

FIG. 9 is also illustrative of the examples hereinbefore given of meansfor automatically controlling engine performance factors, e.g., chamberpressure. In these cases, the sensed parameter is chamber pressure. Thecommand or reference level of chamber pressure is compared with thesensed, i.e., actual chamber pressure, the latter being transduced withthe aid of the interior cowl surface. Any unbalance causes the actuationof the cowl 28, the unbalance constituting the error signal.

In like manner, the control systems of FIG. I and 2 are illustratedgenerally by FIG. 9. The bellows 116, serving as the chamber pressuresensor, and bellows I18 serving to supply the system with the commandedor desired chamber pressure are represented by 13 and 12 in FIG. 9. Thecomparator of that figure comprises the control valve 103 or the switch144, each of which develops an error signal which is fed to therespective actuator 101 and 155 in FIGS. I and 2, and 10 in FIG. 9.

In the case of missile kinematics, the control system reflectsarrangements such as those shown in FIGS. 5 and 6. The accelerometermass or velocity sensing system, represented by parameter sensor 13,supplies one input to the comparator 11 of FIG. 9. The command orreference level is established as described hereinbefore, and isrepresented at 12. An error or correction signal from the comparator lldrives the actuators.

From the foregoing, it may be seen that complete control over missileposition, kinematics and engine parameters is attainable.

D. APPLICATIONS TO OTHER NOZZLE CONFIGURATIONS The control techniqueshereinbefore described may also be utilized with other nozzle types suchas that shown in applicant's above-cited U.S. Pat. Nos. 3, I 5 l ,446and 3,094,072.

E. APPLICATIONS TO LIQUID PROPELLANT ROCKET ENGINES Emphasis has beenplaced upon improvement in solid propellant rocket engines because theflexibility of operation described herein represents a major advance inthe state of the art with respect to conventional solid propellantrocket engines.

Similar improvements may be made in liquid propellant rocket engines aswell. The use of variable throat area and variable expansion rationozzles by application of the principles herein described also permitsgreater flexibility in operation compared to conventional liquidpropellant rocket engines.

Important differences exist in the application of the variable throatarea noule to the two types of rocket engines. In the solid propellantrocket engine, all controls are applied directly to the variable areanozzle, the mass flow rate being a dependent variable whichautomatically increases or decreases in response to changes in throatarea through the influence of chamber pressure on burning rate for agiven propellant grain.

In the liquid propellant rocket engine, the controls may be applieddirectly to the propellant feed system controlling the mass flow rate tothe thrust chamber, the nozzle throat area becoming the dependentvariable which automatically increases or decreases responsive to changein the propellant mass flow rate.

This system is illustrated in FIG. 10 which shows a typical missilepowered by a liquid propellant rocket engine incorporating these newfeatures. For simplicity a liquid monopropellant system is illustrated,with an uncooled thrust chamber. Extension of the design to provide forthe usual bi-propellant system with a regeneratively cooled thrustchamber would follow normal conventional practice to a great degree.

In FIG. 10, 251 is the warhead, and 252 is a guidance system supportedby the airframe 253. Propellant tank 320 contains a liquidmonopropellant which flows through the outlet 321 to the turbo-pumpassembly 322, and is discharged at high pressure through the rigidtubing 323 to a regulating valve 324 mounted on the forward end of thethrust chamber assembly 325. The latter is of novel design having anintegral plug 332 at its aft end and contains radial ports 326 enclosedby a flexibly mounted cowl 327 attached to the chamber by means offlexible connection such as the bellows of FIGS. 2-5 of applicants US.Pat. No. 3,489,373 or a flexible seal, which is insulated as shown inthese figures. The cowl may be positioned by mechanical springs 57 ashereinbefore described, or by fluid springs, such as of FIGS. 3 and 4.

The thrust chamber is rigidly mounted to the thrust structure 328,eliminating the heavy gimbal structure and flexible feed lines otherwiseneeded to oscillate the complete thrust chamber for vector control. Thelatter is achieved instead by rotation of cowl 327 as shown anddescribed hereinbefore.

A small conventional nozzle 329 is formed at the apex of the plugsurface as an additional safety device to prevent accumulation ofunburnt propellant during the starting transient as ignition occurs, thechamber being self-draining.

Variable thrust is controlled by the actuator 330 which operates theworm-gear mechanism 331 to vary the mass flow rate through theregulating valve 324, a butterfly valve being the typical means forvarying the flow rate. The actuator 330 may be a rotary actuator drivingthe worm-gear, or may be replaced by a linear actuator through asuitable lever and linkage system as desired. The actuator may be fluidof electrically driven, being responsive to actions of the variouscontrol systems previously described such as those of FIGS. 1-9.

The operation of the engine may be best described by a simple examplecomparing it with a conventional liquid engine with a constant throatarea nozzle. In the latter case, throttleability is accomplished bycontrol of the propellant feed system, such as by the regulating valveor equivalent means, the reduced mass flow rate reducing chamberpressure with a constant throat area and, hence, thrust. The maximumthrust is then limited by the maximum pumping capacity available, whilethe minimum thrust is limited by the effect of minimum pressure oncombustion stability. For a typical case, the ratio of maximum tominimum pressure may be roughly 3 to l, the thrust ratio differing fromthis slightly due to change in nozzle performance at the two chamberpressures.

With the variable throat area nozzle in FIG. 10, the chamber pressuremay be varied by the propellant feed system in the same manner as in theconstant throat area engine. in this case, however, the nozzle throatarea also increases with increase in chamber pressure. For minimumthrust, the low chamber pressure acting on the differential cowl areacauses only a small cowl extension, or minimum throat area, the productof the two providing minimum thrust. As chamber pressure is increased toits maximum value, the cowl pressure forces further compress the spring57 (or its equivalent fluid spring 160), increasing throat area tomaximum value, the product of the two now providing maximum thrust. Themaximum to minimum thrust ratio is now the product of the maximum tominimum pressure ratio and the maximum to minimum throat area rationeglecting differences in nozzle thrust coefficient as function ofchamber pressure. If the same pressure ratio of 3 now results in, say, afourfold increase in throat area, the thrust ratio for the variable areanozzle becomes approximately 12, instead of 3 with the constant areanozzle. The above numerical values are only illustrative, the actualmaximum to minimum ratios for pressure and area being subject tospecific detail designs.

It may be seen that, for a variable thrust engine, a relatively highspring rate would be desired for the spring 57 (or its equivalent fluidspring 160), for positioning the cowl in order to provide a substantialpressure change for the corresponding area change. This is in contrastto the system described for maintaining constant thrust for the solidpropellant rocket engine under variable ambient temperatures asdescribed previously.

For ballistic missile performance, a variable nozzle expansion ratio asa function of altitude may be provided as described hereinafter forsolid propellant rocket engines. The regressive thrust-timecharacteristic of the solid propellant may be simulated in the liquidrocket installation of FIG. by adapting the actuator, 330, to causegradual closing of the regulating valve 324 as a function of time. Thismay be done by a constant-speed motor drive for the worm-gear 331, orany of a number of types of control, such as a cam and lever system toprovide any desired valve opening, or mass flow rate, for thrustprogramming as a function of time. In this manner, the cowl willautomatically retract as a function of the reduced flow rate, or reducedchamber pressure, making possible automatic variation of expansion ratiowith altitude as a function of the common parameter, time.

The variable area nozzle also offers improved control over the startingtransient, since ignition can be accom plished at the minimum thrustlevel in very large engines, with similar improvement during theshutdown transient.

Other means for varying the mass flow rate or chamber pressure may beemployed. The speed of the turbo-pump assembly may be controlled, suchas by a regulating valve and actuator, similar to 324 and 330, in thefeed lines to the gas generator driving the turbopump (not shown); or bymultiple injector cavities employing individual on-off valves to eachcavity, thereby controlling the number of injector orifices in use.Equivalent means such as a variable orifice injector in the thrustchamber (not shown) may also be used as well as various combinations ofthe foregoing.

Obviously, manual control may be provided, such as in piloted aircraft,with a single throttle control operating the propellant feed systemonly, while the nozzle throat area is varied automatically, asdescribed.

Further improvement in liquid propellant rocket engines may be gainedwith the plug nozzle in combination with systems employing gaspressurization in place of turbo-pump fed systems. Such rocket engines,sometimes referred to as liquid boosters, sometimes, use a small solidpropellant charge as convenient means for pressurizing the liquid, thegaseous products of combustion generated by the solid propellantexpelling the liquid from a tank into the thrust chamber. Therefore, andminimum The flow rate of the liquid propellant is dependent on the rateof gas generation of the solid propellant and, hence, the pressure,thrust and burning rate of the liquid booster experiences a variation asa result of the temperature sensitivity of the solid propellant charge.Therefore, the variable throat area of the plug nozzle will provideimprovements, such as constant pressure and minimum thrust variationwith a liquid booster similar to that already described for solidpropellant rocket engines.

The plug nozzle will, as in the case of solid propellant booster rocketengines, also provide simplified thrust vector control, therebyeliminating large aerodynamic surfaces for stability which cause highaerodynamic drag and heavy structural loads requiring expensiveattachment fittings. Also eliminated is the need for high precisionnozzle thrust alignment; better performance is provided by reducingdispersion.

A typical liquid propellant gas pressurized liquid booster rocketutilizing the plug nozzle is illustrated in FIG. 11. While theillustration is based on the use of the newer liquid monopropellants,its extension to a bipropellant system is straightforward. ln the designshown, the thrust chamber including the plug surface is regenerativelycooled. For simplicity, the cowl is made of uncooled high temperatureresistant refractory material such as used frequently in uncoolednozzles for solid propellant rocket engines. The cowl may also be liquidcooled if desired by suitable means (not shown) such as regenerativecooling by injection of additional propellant into the annular chamberenclosed by the cowl, or by film-cooling.

In FIG. 11, the monopropellant 339, stored within the pressure vessel ortank, 340, is pressurized by the solid propellant charge 341 within thecontainer 342 when the igniter 343 is energized. An expellent bag, 344,separates the gaseous products of combustion generated by the solidpropellant from the liquid propellant preventing chemical interaction,aeration of the liquid, and heat transfer between the hot gases and coolliquid, which have the efiect of producing erratic pressures as thegases are cooled. The expellent bag is in the form of a sleeve seal, oneend of which has a bead, 345, which engages a groove, 346, in the headclosure, 347. The latter is structurally attached by the key, 348, tothe flange, 349, of the tank, 340. The bead 345 then forms a seal,similar to an O-ring seal, between the head closure and tank, and alsoretains the sleeve within the tank. The opposite end of the sleeve,which folds back on itself, is permanently bonded at 350 to the flexiblediaphragm 351. A burst disc, 352, at the lowest portion of the aft headclosure 353 of the tank 340 seals the liquid propellant within the tank340.

The thrust chamber assembly, 360, is fabricated as a separatesub-assembly to facilitate production and attached to the aft head 353of tank 340 by bolts, 354, through the flange, 355, mating with theboss, 356, on the head, 353, although a welded connection may beoptional. The thrust chamber assembly comprises an outer cylinder 357 towhich is attached a plug surface, 358, forming an annular gap, 359, withrespect to an inner cylinder 361 and an attached inner plug surface 362,the gap thus formed providing a coolant passage for the regenerativelycooled thrust chamber. A thrust chamber head closure, 363, completes thecoolant passage by providing a gap in relation to the tank aft closure353.

The injector comprises simple orifices, 364, through the inner plugsurface near the apex of the inner plug, the emerging jet impinging onthe igniter rod, 365, which is inserted through the opening, 366, in theouter plug surface, 358, and retained therein such as by the thread,367. Small bleed holes, 371, may be provided through the outer plugsurface, 358, near the apex to provide local film cooling of the plugtip as required.

The igniter surface is coated with a catalytic material, such as lithiumor its compounds, which reacts with the incoming monopropellant to causehypergolic ignition.

Radial ports, 368, are provided through the inner and outer cylinderwalls of the thrust chamber by the collar, 369, for passage ofcombustion products radially outward into the annular volume enclosed bythe cowl, 370, in the manner previously described. The cowl, 370, isflexibly mounted to the cylinder, 357, such as by a bellows, or flexibleseal, illustrated and described in greater detail hereinbefore. Meansfor positioning the cowl, such as mechanical or fluid springs, oractuators such as 203, including means for thrust vector control, mayalso be as previously described. The various controls for variablethrust may also be used, with the knowledge that there will be a greatertime lag due to the large volume of stored gas as burning timeprogresses.

Means for thrust termination may be provided as part of the cowlassembly, as previously described, since burst discs provide no meansfor terminating propellant flow. A control valve may replace the burstdisc when thrust termination is required, although burst discs mayprovide greater reliability and freedom from leakage during prolongedstorage.

In operation, when the igniter, 343, is energized, the solid propellantcharge generates gas pressure which acts on the diaphragm, 351, causingthe expellent bag to extend downwardly toward the tank aft head, 353.Upon pressurization, the burst disc 352 fails, permitting propellant toenter the coolant passage 359 whence it is injected into the thrustchamber through orifices, 364, striking the igniter, 365, therebyproducing positive ignition. After ignition is initiated, combustion maycontinue by the presence of hot gases within the chamber, with nofurther need for the igniter.

As the propellant level reduces near the end of burning, the flexiblediaphragm 351 will deflect to conform to the shape of the aft headclosure 353 thereby assuring maximum volumetric efficiency approachingl00 percent displacement of stored propellant. Since cer tainmonopropellants have, in general, the properties of Nitric Acid, thesleeve and diaphragm of the expellent bag may preferably be made of acidresistant plastic materials, such as the newer fluorelastomers, whichexhibit good chemical resistance, and good mechanical properties over awide temperature range.

An alternate means for gas pressurization of liquid propellant rocketengines is shown in FIG. 12, in which the simple container 342 of FIG.11 is replaced by a pressure vessel, 342, having a controllable variablearea discharge orifice at 431. The assembly including the pressurevessel 342', solid propellant charge 341, variable area orifice 431becomes a gas generator capable of having a variable and controllablemass flow rate.

Following ignition of the solid propellant by energizing the igniter 343electrically, the internal pressure within the pressure vessel 342' iscontrolled independently of the pressure which may exist at any time inthe tank 340 by controlling the discharge orifice area, 431. The latteris controlled by the actuator 432, supported by bracket 436, by means ofthe rod, 433, which varies the position of the valve, 434, relative tothe seat, 435. As the orifice area is varied, gas generator pressurevaries, and through the dependence of propellant burning rate onpressure, the gas mass flow rate varies. This, in turn, varies the massflow rate of the liquid propellant providing variable thrust. In thisarrangement, simple mechanical or fluid springs may be used to positionthe cowl, 370, of the thrust chamber assembly, 360, of FIG. 11, thevariable thrust being controlled by the solid propellant gas generatorwith variable mass flow rate.

The actuator, 432, controlling the rate of gas generation, may also bedriven by a timing device, similar to that described for the actuator,330, in the turbo-pump fed system of FIG. 10; the gas pressurized liquidbooster rocket engine may then also have the regressive thrust-timecharacteristic similar to the turbo pump system, and analogous toregressive burning solid propellant rocket engines described previouslyand in further detail hereinafter. Either of the three systems may thenbe used for various missile applications such as those hereinafterdescribed.

The variable mass flow gas generator may also automatically compensatefor variation in internal ballistics of the liquid propellant asfunction of ambient temperature and other variables as previouslydescribed for solid propellant rockets.

Thus, it may be seen that new concepts in thrust chamber design,variable mass flow gas generators, simplified injector, expellent bag,general assembly and manufacture, make possible improved performance forliquid propellant booster rocket engines.

For applications to multiple stage rocket engines particularly longrange ballistic missiles, reference is directed to applicantscodivisional application, Ser. No. 784,818 for Solid Propellant RocketEngine Control And Missile Configurations, hereinbefore cited;

What is claimed is:

1. A missile system comprising a missile structure having a store ofpropellant,

a thrust-producing engine for propelling said missile,

said engine having a variable controllable cowl-plug structureconfigured and arranged to affect the throat-area geometry of the engineexhaust structure in a manner to effect continuous control over thethrust producing exhaust therefrom for missile control, actuator meanscoupled to said engine for regulating the configuration of saidcowl-plug structure, and control means operatively coupled to saidactuator means for controlling missile flight, said control meansincluding transducer means for sensing a missile flight parameter to becontrolled,

reference means for establishing a reference signal corresponding to thedesired condition of said parameter,

comparator means responsive to said reference means and transducermeans, and operatively coupled to said actuator means for varying theconfiguration of said engine cowl-plug structure as a function of therelation between said sensed missile flight parameter and said desiredreference parameter condition.

2. A missile system as defined in claim 1, in which said actuator meansinclude means for varying the configuration of said cowl-plug structureto control the stream direction of said thrust producing engine exhaustto thereby vary the attitude of said missile, and said transducer meansinclude means for sensing missile attitude.

3. A missile system as defined in claim 2 in which said reference meansinclude means for supplying a variable command attitude signal to saidcomparator means.

4. A missile as defined in claim 2 in which said reference means includemeans for establishing a reference missile attitude signal for saidcomparator means.

5. A missile system as defined in claim 1 in which said transducer meanscomprises means for sensing the operative pressure within said engineand force prod ucing means operatively coupled to said pressure sensingmeans for controlling the configuration of said cowlplug structure as afunction of said engine pressure. 5 6. A missile system as defined inclaim 1 in which said comparator means comprise signal switching means.

7. A missile system as defined in claim 6 in which said signal switchingmeans comprise electrical switch means.

8. A missile system as defined in claim 6 in which said signal switchingmeans comprise fluid pressure switching means.

9. A missile system as defined in claim 1 in which said transducer meansand reference means comprise fluid pressure translating means.

10. A missile system as defined in claim 1 in which said reference meanscomprise fluid pressure responsive translating means and controllablepressurization means coupled to said translating means for supplyingfluid to said translating means at a pressure corresponding to thedesired condition for said parameter.

11. A missile system as defined in claim 1 in which said actuator meansinclude means for varying the angular relationship between said cow] andplug.

12. A missile system as defined in claim 1 in which said actuator meansinclude means for varying the longitudinal relationship between saidcowl and plug.

13. A missile system as defined in claim I in which said actuator meansinclude means for varying the longitudinal relationship and angularrelationship between said cowl and plug.

14. A missile system comprising a missile structure 35 having a store ofpropellant,

a thrust-producing engine for propelling said missile,

said engine having a variable controllable cowl-plug structureconfigured and arranged to affect the throat-area geometry of the engineexhaust structure in a manner to effect continuous control over thethrust producing exhaust therefrom for missile control,

actuator means coupled to said engine for regulating the configurationof said cowl-plug structure,

and control means operatively coupled to said actuator means forcontrolling missile flight, said control means 50 including transducermeans for sensing a missile flight parameter to be controlled,

reference means for establishing a reference signal corresponding to thedesired condition of said parameter,

comparator means responsive to said reference means and transducermeans, and operatively coupled to said actuator means for varying theconfiguration of said engine cowl-plug structure as a function of therelation between said sensed missile flight parameter and said desiredreference parameter condition, said transducer means including means forsensing the course angle of said missile in flight and generating avariable signal in response to said course angle and said control meansincluding computer means operatively associated with said referencemeans for converting said course angle signal into a missile attitudecontrol signal.

15. A missile system comprising a missile structure having a store ofpropellant,

a thrust-producing engine for propelling said missile,

said engine having a variable controllable cowl-plug structure configured and arranged to affect the throat-area geometry of the engineexhaust structure in a manner to effect continuous control over thethrust producing exhaust therefrom for missile control,

actuator means coupled to said engine for regulating the configurationof said cowl-plug structure, and control means operatively coupled tosaid actuator means for controlling missile flight, said control meansincluding transducer means for sensing a missile flight parameter to becontrolled,

reference means for establishing a reference signal corresponding to thedesired condition of said parameter,

comparator means responsive to said reference means and transducermeans, and operatively coupled to said actuator means for varying theconfiguration of said engine cowl-plug structure as a function of therelation between said sensed missile flight parameter and said desiredreference parameter condition, said transducer means comprising meansfor sensing missile acceleration to control same.

16. A missile system as defined in claim in which said reference meansinclude means for supplying said comparator means with a signalcorresponding to desired missile acceleration.

17. A missile system as defined in claim 16 in which said control meansinclude timing means for regulating the period of said acceleration.

18. A missile system as defined in claim 1 in which said transducermeans comprise spring-mass means.

19. A missile system as defined in claim 1 in which said transducermeans comprise means for sensing missile velocity to control same.

20. A missile system as defined in claim 19 in which said referencemeans comprise means for supplying said comparator means with a signalcorresponding to desired missile velocity.

21. A missile system as defined in claim 1 in which said reference meanscomprise programmed means for varying said desired conditions.

22. A missile system as defined in claim 1 in which said transducermeans include missile acceleration sensing means and missile velocitysensing means.

23. A missile system comprising a missile structure having a store ofpropellant,

a thrust-producing engine for propelling said missile,

said engine having a variable controllable cowl-plug structureconfigured and arranged to affect the throat-area geometry of the engineexhaust structure in a manner to effect continuous control over thethrust producing exhaust therefrom for missile control,

actuator means coupled to said engine for regulating the configurationof said cowl-plug structure, and control means operatively coupled tosaid actuator means for controlling missile flight, said control means 5including transducer means for sensing a missile flight parameter to becontrolled,

reference means for establishing a reference signal corresponding to thedesired condition of said parameter,

comparator means responsive to said reference means and transducermeans, and operatively coupled to said actuator means for varying theconfiguration of said engine cowl-plug structure as a function of therelation between said sensed missile flight parameter and said desiredreference parameter condition, said transducer means including missileacceleration sensing means and missile velocity sensing means and saidcontrol means including means for selectively connecting either saidacceleration sensing means or said velocity sensing means into saidcontrol means.

24. A missile system comprising a missile structure having a store ofpropellant,

a thrust-producing engine for propelling said missile, said enginehaving a variable controllable cowl-plug structure configured andarranged to affect the throat-area geometry of the engine exhauststructure in a manner to effect continuous control over the thrustproducing exhaust therefrom for missile control,

actuator means coupled to said engine for regulating the configurationof said cowl-plug structure,

and control means operatively coupled to said actuator means forcontrolling missile flight, said control means including transducermeans for sensing a missile flight parameter to be controlled,

reference means for establishing a reference signal corresponding to thedesired condition of said parameter,

comparator means responsive to said reference means and transducermeans, and operatively coupled to said actuator means for varying theconfiguration of said engine cowl-plug structure as a function of therelation between said sensed missile flight parameter and said desiredreference parameter condition, said control means further includingmeans for establishing a non-propulsive storage condition for saidsystem.

25. A missile system as defined in claim 24 in which said non-propulsivestorage means include means for adjusting said reference means to astate producing a nonpropulsive configuration of said cowl-plugstructure.

26. A missile system as defined in claim 24 in which said nonpropulsivestorage means comprise means for rendering said actuator meansinoperative to regulate said cowl-plug structure and wherein saidcowl-plug structure normally independently assumes a nonpropulsiveconfiguration in response to internal engine pressure whereby accidentalignition drives said cowlplug configuration to said non-propulsiveconfiguration.

27. A missile system comprising a missile structure having a store ofpropellant,

a thrust producing reaction engine in operative communication with saidpropellant store, said engine having means for controllably varying thethrust produced thereby,

sensor means for monitoring kinematic parameters of linear motion ofsaid missile structure during flight,

and control means operatively connected between said sensor means andsaid thrust varying means to control the flight kinematics of saidmissile.

28. A missile system as defined in claim 27 in which said sensor meanscomprise means for monitoring misaims

1. A missile system comprising a missile structure having a store ofpropellant, a thrust-producing engine for propelling said missile, saidengine having a variable controllable cowl-plug structure configured andarranged to affect the throat-area geometry of the engine exhauststructure in a manner to effect continuous control over the thrustproducing exhaust therefrom for missile control, actuator means coupledto said engine for regulating the configuration of said cowl-plugstructure, and control means operatively coupled to Said actuator meansfor controlling missile flight, said control means including transducermeans for sensing a missile flight parameter to be controlled, referencemeans for establishing a reference signal corresponding to the desiredcondition of said parameter, comparator means responsive to saidreference means and transducer means, and operatively coupled to saidactuator means for varying the configuration of said engine cowl-plugstructure as a function of the relation between said sensed missileflight parameter and said desired reference parameter condition.
 2. Amissile system as defined in claim 1, in which said actuator meansinclude means for varying the configuration of said cowl-plug structureto control the stream direction of said thrust producing engine exhaustto thereby vary the attitude of said missile, and said transducer meansinclude means for sensing missile attitude.
 3. A missile system asdefined in claim 2 in which said reference means include means forsupplying a variable command attitude signal to said comparator means.4. A missile as defined in claim 2 in which said reference means includemeans for establishing a reference missile attitude signal for saidcomparator means.
 5. A missile system as defined in claim 1 in whichsaid transducer means comprises means for sensing the operative pressurewithin said engine and force producing means operatively coupled to saidpressure sensing means for controlling the configuration of saidcowl-plug structure as a function of said engine pressure.
 6. A missilesystem as defined in claim 1 in which said comparator means comprisesignal switching means.
 7. A missile system as defined in claim 6 inwhich said signal switching means comprise electrical switch means.
 8. Amissile system as defined in claim 6 in which said signal switchingmeans comprise fluid pressure switching means.
 9. A missile system asdefined in claim 1 in which said transducer means and reference meanscomprise fluid pressure translating means.
 10. A missile system asdefined in claim 1 in which said reference means comprise fluid pressureresponsive translating means and controllable pressurization meanscoupled to said translating means for supplying fluid to saidtranslating means at a pressure corresponding to the desired conditionfor said parameter.
 11. A missile system as defined in claim 1 in whichsaid actuator means include means for varying the angular relationshipbetween said cowl and plug.
 12. A missile system as defined in claim 1in which said actuator means include means for varying the longitudinalrelationship between said cowl and plug.
 13. A missile system as definedin claim 1 in which said actuator means include means for varying thelongitudinal relationship and angular relationship between said cowl andplug.
 14. A missile system comprising a missile structure having a storeof propellant, a thrust-producing engine for propelling said missile,said engine having a variable controllable cowl-plug structureconfigured and arranged to affect the throat-area geometry of the engineexhaust structure in a manner to effect continuous control over thethrust producing exhaust therefrom for missile control, actuator meanscoupled to said engine for regulating the configuration of saidcowl-plug structure, and control means operatively coupled to saidactuator means for controlling missile flight, said control meansincluding transducer means for sensing a missile flight parameter to becontrolled, reference means for establishing a reference signalcorresponding to the desired condition of said parameter, comparatormeans responsive to said reference means and transducer means, andoperatively coupled to said actuator means for varying the configurationof said engine cowl-plug structure as a function of the relation betweensaid sensed missile flight parameter and said desired referenceparameter condition, said tranSducer means including means for sensingthe course angle of said missile in flight and generating a variablesignal in response to said course angle and said control means includingcomputer means operatively associated with said reference means forconverting said course angle signal into a missile attitude controlsignal.
 15. A missile system comprising a missile structure having astore of propellant, a thrust-producing engine for propelling saidmissile, said engine having a variable controllable cowl-plug structureconfigured and arranged to affect the throat-area geometry of the engineexhaust structure in a manner to effect continuous control over thethrust producing exhaust therefrom for missile control, actuator meanscoupled to said engine for regulating the configuration of saidcowl-plug structure, and control means operatively coupled to saidactuator means for controlling missile flight, said control meansincluding transducer means for sensing a missile flight parameter to becontrolled, reference means for establishing a reference signalcorresponding to the desired condition of said parameter, comparatormeans responsive to said reference means and transducer means, andoperatively coupled to said actuator means for varying the configurationof said engine cowl-plug structure as a function of the relation betweensaid sensed missile flight parameter and said desired referenceparameter condition, said transducer means comprising means for sensingmissile acceleration to control same.
 16. A missile system as defined inclaim 15 in which said reference means include means for supplying saidcomparator means with a signal corresponding to desired missileacceleration.
 17. A missile system as defined in claim 16 in which saidcontrol means include timing means for regulating the period of saidacceleration.
 18. A missile system as defined in claim 1 in which saidtransducer means comprise spring-mass means.
 19. A missile system asdefined in claim 1 in which said transducer means comprise means forsensing missile velocity to control same.
 20. A missile system asdefined in claim 19 in which said reference means comprise means forsupplying said comparator means with a signal corresponding to desiredmissile velocity.
 21. A missile system as defined in claim 1 in whichsaid reference means comprise programmed means for varying said desiredconditions.
 22. A missile system as defined in claim 1 in which saidtransducer means include missile acceleration sensing means and missilevelocity sensing means.
 23. A missile system comprising a missilestructure having a store of propellant, a thrust-producing engine forpropelling said missile, said engine having a variable controllablecowl-plug structure configured and arranged to affect the throat-areageometry of the engine exhaust structure in a manner to effectcontinuous control over the thrust producing exhaust therefrom formissile control, actuator means coupled to said engine for regulatingthe configuration of said cowl-plug structure, and control meansoperatively coupled to said actuator means for controlling missileflight, said control means including transducer means for sensing amissile flight parameter to be controlled, reference means forestablishing a reference signal corresponding to the desired conditionof said parameter, comparator means responsive to said reference meansand transducer means, and operatively coupled to said actuator means forvarying the configuration of said engine cowl-plug structure as afunction of the relation between said sensed missile flight parameterand said desired reference parameter condition, said transducer meansincluding missile acceleration sensing means and missile velocitysensing means and said control means including means for selectivelyconnecting either said acceleration sensing means or said velocitysensing means into said controL means.
 24. A missile system comprising amissile structure having a store of propellant, a thrust-producingengine for propelling said missile, said engine having a variablecontrollable cowl-plug structure configured and arranged to affect thethroat-area geometry of the engine exhaust structure in a manner toeffect continuous control over the thrust producing exhaust therefromfor missile control, actuator means coupled to said engine forregulating the configuration of said cowl-plug structure, and controlmeans operatively coupled to said actuator means for controlling missileflight, said control means including transducer means for sensing amissile flight parameter to be controlled, reference means forestablishing a reference signal corresponding to the desired conditionof said parameter, comparator means responsive to said reference meansand transducer means, and operatively coupled to said actuator means forvarying the configuration of said engine cowl-plug structure as afunction of the relation between said sensed missile flight parameterand said desired reference parameter condition, said control meansfurther including means for establishing a non-propulsive storagecondition for said system.
 25. A missile system as defined in claim 24in which said non-propulsive storage means include means for adjustingsaid reference means to a state producing a non-propulsive configurationof said cowl-plug structure.
 26. A missile system as defined in claim 24in which said non-propulsive storage means comprise means for renderingsaid actuator means inoperative to regulate said cowl-plug structure andwherein said cowl-plug structure normally independently assumes anon-propulsive configuration in response to internal engine pressurewhereby accidental ignition drives said cowl-plug configuration to saidnon-propulsive configuration.
 27. A missile system comprising a missilestructure having a store of propellant, a thrust producing reactionengine in operative communication with said propellant store, saidengine having means for controllably varying the thrust producedthereby, sensor means for monitoring kinematic parameters of linearmotion of said missile structure during flight, and control meansoperatively connected between said sensor means and said thrust varyingmeans to control the flight kinematics of said missile.
 28. A missilesystem as defined in claim 27 in which said sensor means comprise meansfor monitoring missile velocity.
 29. A missile system as defined inclaim 27 in which said sensor means comprise means for monitoringmissile acceleration.
 30. A missile system as defined in claim 27 inwhich said control means comprise differential fluid valve meansresponsively coupled to said sensor means and hydraulic actuator meansresponsively coupled to said valve means.
 31. A missile system asdefined in claim 27 in which said control means comprise differentialswitch means responsively coupled to said sensor means and abi-directional electrical actuator responsively coupled to said switchmeans.